A conventional gas turbine engine includes a compressor, a combustor and a turbine. The compressor compresses ambient air which is supplied to the combustor where the compressed air is combined with a fuel and ignites the mixture, creating combustion products defining a working gas. The working gas is supplied to the turbine where the gas passes through a plurality of paired rows of stationary vanes and rotating blades. The rotating blades are coupled to a shaft and disc assembly. As the working gas expands through the turbine, the working gas causes the blades, and therefore the shaft and disc assembly, to rotate.
Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine vanes and blades must be made of materials capable of withstanding such high temperatures. In addition, turbine vanes and blades often contain internal cooling systems for prolonging the life of the vanes and blades and reducing the likelihood of failure as a result of excessive temperatures.
Typically, turbine vanes comprise inner and outer endwalls and an airfoil that extends between the inner and outer endwalls. The airfoil is ordinarily composed of pressure and suction sidewalls extending between a leading edge and a trailing edge. The vane cooling system receives air from the compressor of the turbine engine and passes the air through the airfoil. One example of a cooling system within a vane is disclosed in U.S. Pat. No. 6,955,523. The cooling system comprises a cooling circuit formed configured as a serpentine cooling path to effect cooling of the airfoil wall.
Known serpentine cooling systems with low cooling flow rates and a large cross-sectional ratio between the inner and outer endwalls may experience diffusion flow problems and a corresponding decreased heat transfer coefficient. For example, known turbine vane airfoil cooling designs have resolved the diffusion problem for a low mass flux serpentine flow channel by including a by-pass for allowing a portion of the cooling air to flow in between the upstream and downstream serpentine flow channels. The by-pass air facilitates maintaining the through flow channel Mach number, particularly in the large cross-sectional area portions of the vane located toward the outer endwall.
Accordingly, it is desirable to improve the heat transfer characteristics of cooling fluid flowing through turbine airfoils having large cross-sectional ratios between inner and outer ends of the airfoil. In particular, it is desirable to fully utilize the serpentine flow network within an airfoil, such as by avoiding a flow by-pass, including minimizing the adverse affects of diffusion flow by maintaining the Mach number as cooling fluid is conducted throughout the cooling circuit.